This disclosure relates to a gas turbine engine component, such as an airfoil. More particularly, the disclosure relates to a cooling configuration used to effectively turn the cooling fluid at two adjacent cooling fluid exits.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades.
Many blades and vanes, blade outer air seals, turbine platforms, and other components include internal cooling passages. Blade trailing edge tips are typically susceptible to high temperature damage such as thermal mechanical fatigue, and/or oxidation. Current airfoil designs tend to leave the trailing edge tip without convective cooling.
To address this concern, one example airfoil design adds flow features at the tip, which enables the entire trailing edge to be convectively cooled. Discrete slots or a large continuous slot is provided the length of the trailing edge all the way to the tip.